Combined-cycle propulsion is considered when the high efficiency of air-breathing propulsion is desired over a broad Mach number range. Air-breathing access to space is one such application of current interest to NASA. The dual-mode scramjet is central to most combined-cycle schemes. Turbine-based combined-cycle (TBCC) systems use a turbine engine for low speed acceleration, and operate to a maximum flight Mach number in scramjet mode dictated by system considerations. TBCC systems are normally assumed to take-off horizontally, and use a second, rocket-powered stage to achieve orbit. Rocket-based combined-cycle (RBCC) schemes use chemical rocket propulsion for low speed acceleration. The high thrust-to-weight ratio of the rocket component allows for its integration within the air-breathing duct. RBCC systems are normally assumed to be launched vertically, and can operate from lift-off to orbit. Turbine-engines reach temperature and thrust limitations as Mach number increases. Rocket thrusters provide a high ratio of thrust-to-weight at any speed, but are very inefficient from the standpoint of specific impulse. In either case, it is advantageous to extend dual-mode scramjet operation to as low a Mach number as possible.
Supersonic combustion has long been recognized as a solution to problems associated with the severe stagnation conditions within a ramjet engine at high flight Mach number. Diffuser momentum loss, dissociation, non-equilibrium expansion losses, and structural loading are all relieved by transition to a supersonic combustion process. In general, the cross-sectional area of the supersonic combustor increases in the downstream direction to avoid thermal choking and excessive pressure gradients. The subsequent nozzle expansion process requires a more dramatic increase in cross-sectional area and is usually integrated with the vehicle aft end to provide the maximum possible area ratio.
In order to extend the operable flight Mach number range of the scramjet engine downward, toward the upper limit for turbojets or to limit rocket operation to as low a ΔV (speed range) as possible, “dual-mode” operation was introduced by Curran, et al. in U.S. Pat. No. 3,667,233. U.S. Pat. No. 3,667,233 is incorporated herein by reference hereto.
FIG. 1 is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic diagram partially in block form of a dual mode combustion chamber according to the invention.
FIG. 2 is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic cross section of the device of FIG. 1 showing one possible arrangement of the fuel injectors.
FIG. 3 is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic diagram partially in block form showing an annular configuration for the combustion chamber of FIG. 1.
FIG. 4 is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic end view of the device of FIG. 3 from the exhaust end.
FIG. 5 is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic diagram partially in block form of a modified fuel supply system for the device of FIG. 1.
Conceptually, a thermally-choked combustion process is established in the aft regions of the scramjet flowpath where the cross-sectional areas are greatest. As depicted in FIG. 7, the diverging scramjet duct acts as a subsonic diffuser, and the thermal throat is positioned so as to produce the required back-pressure.
FIGS. 6 and 7 are another illustration of the structure and process of the prior art Curran et al., U.S. Pat. No. 3,667,233.
Curran et al., U.S. Pat. No. 3,667,233, states at col. 1, lns. 29 et seq. that:
“A combustor with a fixed geometry has one parallel combustion section with a substantially uniform cross section along its length. Fuel is injected into this section and the flame is stabilized on recessed flameholders. As the fuel burns it causes choked flow in this section which sends a shock wave upstream to convert the normal supersonic flow through the combustor to subsonic flow. For transition from subsonic mode to the supersonic mode, fuel is injected into a diverging section upstream of the parallel section which causes the shock to move downstream until it is ejected from the engine. In the final transition to supersonic mode, fuel is supplied only to the upstream injectors.”
Further, Curran et al., U.S. Pat. No. 3,667,233, states at col. 2, lns. 4 et seq. that:
“At these speeds fuel is supplied to nozzles 36. Burning of the fuel in the uniform cross section combustion chamber 24 causes choked flow which sends a shock wave upstream of the flow to convert the supersonic flow to subsonic flow within the combustion chamber. As the speed of the aircraft increases to a speed between Mach 4 and Mach 5, fuel control 30 starts a flow of fuel to nozzles 32 as the fuel control 34 gradually decreases the fuel flow to nozzles 36. This causes the shock wave to gradually recede as fuel to nozzles 32 is increased and fuel flow is decreased to nozzles 36. At a speed of about Mach 8 fuel to nozzles 36 is further reduced and supersonic combustion now occurs throughout the divergent and parallel ducts. The expansion of the heated gases in expansion section 22 permits higher Mach speeds to be attained.”
The cross-sectional area of the thermal throat must increase as flight Mach number decreases, unless fuel-to-air ratio is reduced. For a given duct, this effect determines the minimum flight Mach number for dual-mode operation. At Mach 2.5, the required thermal throat area approaches that of the inlet capture area. The primary technical challenges in practical application of the dual-mode scramjet scheme of Curran et al., U.S. Pat. No. 3,667,233, are modulation of the thermal throat location, modulation of fuel distribution, ignition, and flame-holding in the large cross-section. Any in-stream devices must be retractable or expendable so as not to inhibit supersonic combustion operation.
Curran et al., U.S. Pat. No. 3,667,233, controls fuel flow to modulate the position of the thermal throat at low flight Mach numbers and then, subsequently, to transition to supersonic ramjet operation. If Curran doesn't modulate the position of the choked flow correctly, the shock wave moves further upstream into the inlet passage 21 of Curran and un-start of the engine may occur.
FIG. 6 shows a cross-sectional view 600 of a prior art (Curran et al.) scramjet engine operating in the scramjet mode. Processes that govern scramjet efficiency are inlet momentum losses, Rayleigh losses due to heat addition, heat loss to the combustor walls, skin friction, and non-equilibrium expansion. Other factors that must be considered include separation of boundary layers due to adverse pressure gradients, intense local heating at re-attachment points and shock impingements, and fuel staging or variable geometry to accommodate the variation of combustion area ratio with free stream stagnation enthalpy.
Referring to FIG. 6, fuel injection nozzle 601, inlet contraction section 602, diverging supersonic combustion section 603, and exit nozzle 604 are illustrated. As stated above, in the scramjet mode, this engine is fed with fuel injector 601. Reference numeral 608 illustrates and internal wall of the engine. Reference numeral 606 signifies incoming air being compressed. Reference numeral 605 represents the fuel-air mixture being combusted. And, reference numeral 607 signifies expanded gas/combustion products being expelled from the engine.
FIG. 7 is the cross-sectional view 700 of FIG. 6 (Curran et al. prior art engine) in the ramjet mode illustrating choked flow 702 and a shock waves 701. Fuel injectors 703, 704 are illustrated and are operable in the ramjet mode. Curran et al. must delicately control the insertion of fuel. First, fuel is inserted with injectors 703, 704 and then fuel is inserted using injector 601 to prevent the shock wave from being expelled leftwardly into the inlet contraction section 602 which may result in un-start of the engine. Reference numeral 606A indicates incoming compressed air and reference numeral 607A represents combustion products expelled from the engine.